1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with tip cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, the turbine section includes a plurality of stages of turbine rotor blades with blade tips that from a gap with an outer shroud of the engine in which the hot gas flow passing through the turbine can leak past. The blade tip gap leakage not only reduces the efficiency of the turbine by not impacting all of the gas flow onto the turbine rotor blades, but can cause thermal damage to the blade tips and result in shortened life for the blades.
In a high temperature turbine blade tip section, the heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. Thus, blade tip section sealing and cooling must be addressed as a single problem. In the prior art, a turbine blade tip includes a squealer tip rail that extends around the perimeter of the airfoil flush with the airfoil wall and forms an inner squealer pocket. The main purpose of using a squealer tip in a blade design is to reduce the blade tip leakage and also to provide the rubbing capability for the blade.
In the prior art, blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine coolant passages from both the pressure and suction surfaces near the blade tip edge and the top surface of the squealer cavity. In general, film cooling holes are located along the airfoil pressure side and suction side tip sections and from the leading edge to the trailing edge to provide edge cooling for the blade squealer tip. In addition, convective cooling holes are also located along the tip rail at the inner portion of the squealer pocket to provide for additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow field, a large quantity of film cooling holes and cooling flow is required in order for adequate cooling of the blade tip periphery.
FIG. 1 shows a prior art rotor blade squealer tip cooling design with the secondary hot gas flow migration around the blade tip section. The squealer tip pocket is formed by the pressure side and the suction side walls and the pocket floor. Film cooling holes are shown on the pressure side wall just beneath the squealer tip edge. Cooling holes are shown on the pocket floor to discharge cooling air from the internal cooling air passage and into the squealer pocket. The airflow over the blade tip flows in a vortex pattern as indicated by the arrows. FIGS. 2 and 3 shows the pressure side film cooling hole arrangement and shape of each film cooling hole opening. FIGS. 4 and 5 shows the suction side film cooling hole arrangement and shape of each film cooling hole opening.
The blade squealer tip rail is subject to heating from three exposed sides which are heat load form the airfoil hot gas side surface of the tip rail, heat load from the top portion of the tip rail, and heat load from the back side of the tip rail. Cooling of the squealer tip rail by means of discharge row of film cooling holes along the blade pressure side and suction side peripheral and conduction through the base region of the squealer tip becomes insufficient. This is primarily due to the combination of squealer pocket geometry and the interaction of hot gas secondary flow mixing. The effectiveness induced by the pressure film cooling and the tip section convective cooling holes becomes very limited.
U.S. Pat. No. 6,494,678 B1 issued to Bunker on Dec. 17, 2002 and entitled FILM COOLED BLADE TIP discloses a turbine rotor blade with a tip having multi-channel cooling grooves (#50 in the Bunker patent) arranged along the tip edge on the pressure side wall of the blade and discharge cooling air from an internal cooling channel of the blade. These cooling holes and channels do not collect and accelerate the cooling air as in the present invention.
It is therefore an object of the present invention to provide for a turbine rotor blade with film cooling holes for the blade tip edge periphery that greatly reduces the airfoil tip edge metal temperature and therefore reduces the cooling flow requirement and improve turbine efficiency.
It is another object of the present invention to provide for a turbine rotor blade with film cooling holes for the blade tip edge periphery that will provide for a film of cooling air to pass over the blade tip.
It is another object of the present invention to provide for a turbine rotor blade with notches on the pressure side tip edge that will retain the cooling air longer in the notches.
It is another object of the present invention to provide for a turbine rotor blade with increased tip section cooling side convection wedded surface area.
It is another object of the present invention to provide for a turbine rotor blade with a blade tip with reduced hot side convective area.
It is another object of the present invention to provide for a turbine rotor blade with a blade tip that will reduce the effective leakage flow effective area which reduces leakage flow and thus lowers the heat load onto the blade tip.